Blade outer air seal with passage joined cavities

ABSTRACT

A blade outer air seal assembly includes a body that defines a first cavity separated from a second cavity by a circumferential rib. The circumferential rib includes at least one passage which provides communication between the first cavity and the second cavity.

BACKGROUND

The present application relates to a blade outer air seal (BOAS) and more particularly to a multi-cavity blade outer air seal (BOAS).

Gas turbine engines generally include fan, compressor, combustor and turbine sections along an engine axis of rotation. The fan, compressor, and turbine sections each include a series of stator and rotor blade assemblies. A rotor and an axially adjacent array of stator assemblies may be referred to as a stage. Each stator vane assembly increases efficiency through the direction of core gas flow into or out of the rotor assemblies.

An outer case, including a multiple of blade outer air seals (BOAS), provides an outer radial flow path boundary. A multiple of BOAS are typically provided to accommodate thermal and dynamic variation typical in a high pressure turbine (HPT) section of the gas turbine engine. The BOAS are subjected to relatively high temperatures and receive a secondary cooling airflow for temperature control. The secondary cooling airflow is communicated into the BOAS then through cooling channels within the BOAS for temperature control.

SUMMARY

A blade outer air seal assembly according to an exemplary aspect of the present disclosure includes a body that defines a first cavity separated from a second cavity by a circumferential rib. The circumferential rib includes at least one passage which provides communication between the first cavity and the second cavity.

A method of communicating a secondary cooling airflow within a gas turbine engine according to an exemplary aspect of the present disclosure includes segregating a first cavity from a second cavity by a circumferential rib, the circumferential rib having at least one passage. Communicating secondary cooling airflow between the first cavity and the second cavity through the at least one passage.

BRIEF DESCRIPTION OF THE DRAWINGS

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiment. The drawings that accompany the detailed description can be briefly described as follows:

FIG. 1 is a general sectional diagrammatic view of a gas turbine engine HPT section;

FIG. 2 is a perspective exploded view of a BOAS segment; and

FIG. 3 is a chart of pressures within the BOAS and axial distance from a leading edge thereof;

DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT

FIG. 1 schematically illustrates a gas turbine engine 20, illustrated partially herein as a High Pressure Turbine (HPT) section 22 disposed along a common engine longitudinal axis A. The engine 20 includes a Blade Outer Air Seal (BOAS) assembly 24 to provide an outer core gas path seal for the turbine section 22. It should be understood that although a BOAS assembly for a HPT of a gas turbine engine is disclosed in the illustrated embodiment, the BOAS assembly may be utilized in any section of a gas turbine engine. The BOAS segment may find beneficial use in many industries including aerospace, industrial, electricity generation, naval propulsion, pumping sets for gas and oil transmission, aircraft propulsion, vehicle engines, and stationary power plants.

The HPT section 22 generally includes a rotor assembly 26 disposed between forward and aft stationary vane assemblies 28, 30. Outer vane supports 28A, 30A attach the respective vane assemblies to an engine case 32 (illustrated schematically). The rotor assembly 26 generally includes a multiple of airfoils 34 circumferentially disposed around a disk 36. The distal end of each airfoil 34 may be referred to as an airfoil tip 34T which rides adjacent to the BOAS assembly 24.

The BOAS assembly 24 is generally disposed in an annulus radially between the engine case 32 and the airfoil tips 34T. The BOAS assembly 24 generally includes a blade outer air seal (BOAS) support 38 and a multiple of blade outer air seal (BOAS) segments 40 mountable thereto (also see FIG. 2). The BOAS support 38 is mounted within the engine case 32 to define forward and aft flanges 42, 44 to receive the BOAS segments 40. The forward flanges 42 and the aft flanges 44 may be circumferentially segmented to receive the BOAS segments 40 in a circumferentially rotated and locked arrangement as generally understood. It should be understood that various interfaces and BOAS assemblies may alternatively be provided.

Each BOAS segment 40 includes a body 46 which defines a forward interface 48 and an aft interface 50. The forward interface 48 and the aft interface 50 respectively engage the flanges 42, 44 to secure each individual BOAS segment 40 thereto.

With reference to FIG. 2, each BOAS segment 40 includes at least two cavities 52A, 52B to receive a secondary cooling airflow S. Each cavity 52A, 52B may be formed through, for example, an investment casting process then closed by a single impingement plate 54.

In the disclosed non-limiting embodiment, the cavity 52A is axially forward of cavity 52B but separated therefrom by a circumferential rib 56. That is, the circumferential rib 56 essentially surrounds the engine longitudinal axis A. Secondary cooling air S flows through the plate 54, impinges in the BOAS cavities 52A, 52B then flows out to the core gaspath flow through a multiple of edge holes 60. The circumferential rib 56 and plate 54 isolates the secondary cooling air allocated to a specific cavity 52A, 52B. It should be understood that various alternative cavity and passageway arrangements may be provided.

The circumferential rib 56 includes at least one passage 58 which provides for secondary cooling air S to flow aftward from the forward cavity 52A to the aft cavity 52B. It should be understood that the term “passage” as utilized herein may include various slots, apertures, openings, holes and paths. In the disclosed non-limiting embodiment, three passages 58 are provided which removes approximately 3.5% of the rib 56. Since the percentage of material removed is minimal and since the removal of the material is from a circumferential member rather than an axial member, minimal, if any, structural impact is experienced by the BOAS segment 40.

The passages 58 allows some of the forward cavity 52A secondary cooling air S to be reused in the aft cavity 52B which results in lower temperatures and relatively lower cooling flow requirements for the BOAS segment 40. The passages 58 also permits at least some reuse of the secondary cooling air S with but the single plate 54 which need not be welded to the rib 56. The single plate 54 facilitates manufacture with minimal brazing filler metal (BFM) and minimizes undesired leakage.

In the disclosed non-limiting embodiment, the secondary cooling air S gaspath pressure within the BOAS segment 40 is lower axially aft of airfoil tips 34T (FIG. 1). The forward cavity 52A thus has a somewhat higher static pressure than the aft cavity 52B (FIG. 3) due to the direction of primary core flow. The higher static pressure in cavity 52A also results in increased axial crossflow heat transfer coefficient (Hc) in the forward cavity 52A which results in, for example, lower temperatures, and, likewise, longer operational life of the BOAS segment 40 as represented in the chart below:

RELATED ART 2-Cavity with 3 2-Cavity Total Flow passages - Total Flow % Wae 0.69% 0.59% Tmet, mx (max BOAS 2159° F. 2151° F. temperature) Nominal/min back 13.2%/5% 12.6% flow margin (BFM)

It should be understood that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, it should be understood that steps may be performed in any order, separated or combined unless otherwise indicated and will still benefit from the present invention.

The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The disclosed embodiments of this invention have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. For that reason the following claims should be studied to determine the true scope and content of this invention. 

1. A blade outer air seal assembly comprising: a body which defines a first cavity separated from a second cavity by a circumferential rib, said circumferential rib includes at least one passage which provides communication between said first cavity and said second cavity.
 2. The blade outer air seal assembly as recited in claim 1, further comprising an impingement plate which encloses said first cavity and said second cavity.
 3. The blade outer air seal assembly as recited in claim 1, wherein said at least one passage comprises three passages.
 4. The blade outer air seal assembly as recited in claim 3, wherein said three passages removes approximately 3.5% of said rib.
 5. The blade outer air seal assembly as recited in claim 1, wherein said at least one passage removes approximately 3.5% of said rib.
 6. The blade outer air seal assembly as recited in claim 1, wherein said first cavity is a forward cavity and said second cavity is axially aft of said first cavity.
 7. The blade outer air seal assembly as recited in claim 6, wherein said at least one passage comprises three passages.
 8. The blade outer air seal assembly as recited in claim 7, wherein said three passages removes approximately 3.5% of said rib.
 10. A method of communicating a secondary cooling airflow within a gas turbine engine comprising: segregating a first cavity from a second cavity by a circumferential rib, the circumferential rib having at least one passage; and communicating secondary cooling airflow between the first cavity and the second cavity through the at least one passage.
 11. The method as recited in claim 10, communicating the secondary cooling airflow into the first cavity and the second cavity through a single impingement plate.
 12. The method as recited in claim 11, communicating the secondary cooling airflow from the first cavity and the second cavity through a multiple of edge holes to a core flow. 